The present invention relates to the field of gas turbine engines, more particularly such gas turbine engines having devices for preventing the ingestion of water into the combustion chamber.
Gas turbine engines, particularly turbojet engines used to power aircraft, typically comprise an axial air compressor; a generally annular diffuser located downstream of the air compressor and bounded by an outer wall and an inner wall, the diffuser receiving compressed air from the compressor and directing it towards a combustion chamber; a generally annular combustion chamber in which the compressed air is mixed with pressurized fuel and ignited; and an axial turbine located downstream of the combustion chamber which drives the air compressor upon being rotated by the exhaust gases from the combustion chamber. As is well known in the art, the air passes in an upstream to downstream direction by sequentially passing through the compressor, the diffuser, the combustion chamber and the turbine.
Such turbojet engines with axial geometry at their front intakes not only receive the air required for operation, but also, depending upon climatic conditions, may also ingest sand and/or water which will deleteriously affect their operation. This is particularly true when an aircraft passes through a storm, or a large cloud such as a cumulus or a cumulo-nimbus cloud. In these events, large amounts of water may enter the engine compressor. If the engine is operating at full power, the water will vaporize and, even if it were to enter the combustion chamber, would be atomized sufficiently so as to not extinguish the ignition within the chamber which, under these conditions, is being supplied with fuel at a high rate.
However, when the aircraft is operating under low power conditions, such as descending during a landing approach, liquid water, in the form of large drops or even sheets of water may pass into the engine and be ingested into the combustion chamber. Under these conditions, this amount of water may extinguish the combustion at one or more of the burners, since the fuel supply is at a comparatively low rate. Ignitors, which are located downstream of the fuel injectors, may also be wetted by the water and become temporarily inoperative. Unless the pilot can direct the aircraft out of the atmospheric conditions, the engine may be stopped completely.
To avoid such potentially catastrophic situations, it is mandatory to reduce the quantities of water entering the combustion chamber, or, at the very least, prevent the moist air from extinguishing the burners and wetting the ignitors.
Various solutions have been proposed to reduce the amount of water in a turbojet engine air flow. U.S. Pat. No. 4,389,227 describes a system for collecting a flow of water on a nose cowl and to split it into droplets. Because of their weight, the droplets are evacuated into the cold flow bypass duct, thereby precluding their intake into the combustion chamber.
U.S. Pat. No. 4,255,174 describes a separator to remove liquid droplets from a gas flow. The guide vanes define a plurality of channels spaced along the entire length of their leading edges, which channels combine at the top of the assembly at their trailing edges. The centrifuged droplets are expelled toward the radially outer zones of a primary air duct.
U.S. Pat. No. 5,044,153 describes a turbojet engine wherein water entering the engine is removed by discharge scoops located immediately downstream of the low pressure compressor.
In these systems, the partial elimination of the water taken by the turbojet engine is carried out in the air-flow far upstream of the combustion chamber. Because of the swirling air motion within the engine, a peripheral film of hot air is enriched with upstream water. This peripheral film is especially water enriched at the diffuser exit opposite the fuel injector air intakes. Because of their high density, the water drops of the peripheral film are moved toward the injectors into the combustion chamber, where the ignition may be extinguished and the ignitors may be wetted.